Rotary wing vehicle

ABSTRACT

A rotary wing vehicle includes a body structure having an elongated tubular backbone or core, and a counter-rotating coaxial rotor system having rotors with each rotor having a separate motor to drive the rotors about a common rotor axis of rotation. The rotor system is used to move the rotary wing vehicle in directional flight.

RELATED APPLICATIONS

This application is a continuation-in-part of U.S. application Ser. No.13/270,872, filed Oct. 11, 2011, which is a divisional of U.S.application Ser. No. 12/872,622, filed Aug. 31, 2010 (now U.S. Pat. No.8,042,763, issued Oct. 25, 2011), which is a divisional of U.S.application Ser. No. 11/105,746, filed Apr. 14, 2005 (now U.S. Pat. No.7,789,341, issued Sep. 7, 2010), all of which are incorporated byreference herein in their entirety. This application also claimspriority under 35 U.S.C. §119(e) to U.S. Provisional Application Ser.No. 61/649,741, filed May 21, 2012 and U.S. Provisional Application Ser.No. 61/799,878, filed Mar. 15, 2013, both of which are incorporated byreference herein in their entirety. This application is also related toU.S. Provisional Application No. 60/562,081, filed Apr. 14, 2004, whichis incorporated by reference herein in its entirety.

BACKGROUND

The present disclosure relates to aerial vehicles and particularly tounmanned aerial vehicles (UAV). More particularly, the presentdisclosure relates to unmanned rotary wing vehicles.

Rotary wing vehicles are used in a variety of applications. Unmannedrotary wing vehicles are often used by the military, law enforcementagencies, and commercial activities for aerial reconnaissanceoperations.

SUMMARY

A rotary wing vehicle, in accordance with the present disclosureincludes a body structure having an elongated tubular backbone or coreand a counter-rotating coaxial rotor system having rotors with eachrotor having a separate motor to drive the rotors about a common rotoraxis of rotation. A power source comprising, for example, a battery,fuel cell, or hybrid gas-electric generator is provided to supplyelectric power to the motors. Power transmission to and between therotor systems is accomplished primarily by means of electrical wiringinstead of mechanical shafting. A modular structure is described whichassists manufacturability.

In illustrative embodiments, a torque tube is provided to transmitmechanical power inside the non-rotating tubular backbone creating amodular mast structure that can be used to support coaxial rotor systemson many types of vehicles.

In illustrative embodiments, a blade pitch control system is locatedbetween the rotor blades. A fixed, non-rotating body shell oraerodynamic fairing may be provided between the upper and lower rotorsto protect the pitch control system and airframe against the elementsand to reduce aerodynamic drag of the aircraft.

In illustrative embodiments, an auxiliary power-pack is provided whichis separable from the vehicle in flight to facilitate, for instance,delivery of the vehicle to a distant location. In another embodiment,the power-pack comprises a payload such as an explosive munition,dipping sonar, hydrophones, or a separable sonobouy module. Whileaspects of the disclosure are applicable to many helicopters, includingfull-sized man carrying helicopters, the current disclosure isespecially well suited for application to autonomous or radio-controlledrotary wing aircraft known as remotely piloted vehicles (RPVs), orunmanned aerial vehicles (UAVs).

Additional features of the present disclosure will become apparent tothose skilled in the art upon consideration of illustrative embodimentsexemplifying the best mode of carrying out the disclosure as presentlyperceived.

BRIEF DESCRIPTIONS OF THE DRAWINGS

The detailed description particularly refers to the accompanying figuresin which:

FIG. 1 is a diagrammatic view of a rotary wing vehicle in accordancewith the present disclosure showing an aircraft including a guidancesystem, and a pair of rotor systems coupled to an airframe comprising anon-rotating structural spine or backbone and carrying a payload;

FIG. 2A is a perspective view of a rotary wing vehicle in accordancewith the present disclosure showing a counter-rotating coaxial rotorsystem in a vertical flight mode;

FIG. 2B is a perspective view of the rotary wing vehicle of FIG. 2Ahaving a counter-rotating coaxial rotor system and a fixed-wing boostermodule in a horizontal flight mode;

FIG. 3 is a side elevation view of the rotary wing vehicle of FIG. 2Ashowing exterior body panels, electrical wiring, and booster sectionremoved for clarity;

FIG. 4 is a side elevation view, with portions broken away, of thevehicle of FIG. 2A showing a counter-rotating coaxial rotor system andan electrical power source;

FIG. 5 is an enlarged perspective view of the vehicle of FIG. 2A, withportions broken away, showing an upper interior section of the vehicleand the counter-rotating coaxial rotor system;

FIG. 6 is an enlarged perspective view of the vehicle of FIG. 2A, withportions broken away, showing a lower interior section of the vehicleand the counter-rotating coaxial rotor system;

FIG. 7A is a perspective view of a core tube or backbone having acircular cross section and a hollow interior channel that is used as aconduit between sections of the vehicle and showing electrical wiringrunning through the hollow interior and entering and exiting at variouspoints;

FIG. 7B is a perspective view of backbone having a generally cruciformcross section with exterior channels running the length of the backbonethat can be used as conduits between sections of the vehicle.

FIG. 8 is an enlarged perspective view of a first ring mount;

FIG. 9 is an exploded perspective view of a second ring mount showingattached linkages and body supports;

FIG. 10 is an enlarged perspective view of a middle interior section ofthe vehicle of FIG. 2A, with portions broken away, showing thecounter-rotating coaxial rotor system;

FIG. 11A is an exploded perspective view of a rotor module having rotorblades with variable cyclic pitch and fixed collective pitch;

FIG. 11B is an exploded perspective view of a rotor module having rotorblades with variable cyclic and variable collective pitch;

FIGS. 12A and 12B are perspective views of a first side and a secondside of a motor mount;

FIGS. 13A and 13B are perspective views of a first side and a secondside of a rotor hub;

FIG. 14 is a sectional view taken along lines 14-14 of FIG. 2B, showingthe rotor module;

FIG. 15 is a side elevation view of the counter-rotating coaxial rotorsystem of FIG. 2A, and a core tube depending from the rotor system;

FIGS. 16A and 16B are exploded perspective views of a single powermodule including several batteries;

FIG. 17 is an orthographic view of the booster module of FIG. 2B showingone wing folded for storage and one wing extended in a flightconfiguration;

FIG. 18A is an orthographic view depicting a rotatory wing vehicle inflight after separation from the booster module;

FIG. 18B is an orthographic view depicting the booster module afterseparation from the rotary wing vehicle of FIG. 18A;

FIG. 19 is an elevation view of the rotary wing vehicle showing adipping sonar or hydrophone assembly depending from a bottom portion ofthe vehicle;

FIGS. 20A, 20B, and 20C are sequential views of the rotary wing vehicleshowing the operation of unequal length folding blades during a crashlanding of the vehicle on ground underlying the rotary wing vehicle;

FIGS. 21A and 21B are side elevation views of a storage tube and therotary wing vehicle showing the vehicle folded for storage;

FIG. 22 is a perspective view of a rotary wing vehicle in accordancewith present disclosure delivering a sensor or marking to a remotelocation shown for the purpose of illustration to be a ship on the openocean;

FIG. 23 is a side elevation view of a rotary wing vehicle folded forstorage in a rear portion of a gravity-delivered bomb;

FIG. 24 is a perspective view of a rotary wing vehicle deploying fromthe rear of a gravity-delivered bomb to the vicinity of a target siteshowing the gravity-delivered bomb ejecting the rotary wing vehicle andthe rotary wing vehicle deploying into a vertical flight mode to loiterin the target area to provide an attacking force with real-time battledamage assessment after the gravity delivered bomb has struck thetarget;

FIG. 25A is a diagrammatic view of another rotary wing vehicle showingan aircraft having a central buss architecture with power and signalconduits, a guidance system, and a pair of rotor systems coupled to anairframe comprising a non-rotating structural spine or backbone andcarrying a payload;

FIG. 25 B is a diagrammatic view of the rotary wing vehicle of FIG. 25Ashowing a rotor system, control system, and power supply communicatingthrough a central data/power buss with power and signal conduit;

FIG. 26 is a diagrammatic view of another embodiment of a rotary wingvehicle, according to the present disclosure, having a central bussarchitecture with power and signal conduits, a guidance system, and apair of rotor systems coupled to an air frame;

FIG. 27 is an elevation view of a rotary wing vehicle according to thepresent disclosure showing the rotary wing vehicle includes astreamlined body suited to high-speed translational flight and a coaxialmast module that includes an internal torque tube for driving an upperrotor;

FIG. 28 is an elevation view of the rotary wing vehicle of FIG. 27 withportions of the body shells broken away to reveal the mast module androtor control systems;

FIG. 29A is an enlarged side elevation view of the rotary wing vehicleof FIG. 28 with portions of the mast module and rotor shroud cut away toreveal interior detail;

FIG. 29B is an enlarged portion taken from the circled region of FIG.29A;

FIG. 30 is an elevation view of another embodiment of a rotary wingvehicle in accordance with the present disclosure showing the rotarywing vehicle includes a streamlined body suited to high-speedtranslational flight and a coaxial mast module that includes an upperrotor speed reducer and showing that portions of body shells included inthe streamlined body have been broken away to reveal a mast module androtor control systems;

FIG. 31 is an enlarged elevation view of the rotary wing vehicle of FIG.29 with portions of the mast module and rotor shroud broken away toreveal interior detail;

FIG. 32 is a sectional view of the mast tube of the rotary wing aircraftof FIG. 28;

FIG. 33 is an enlarged perspective view of a servo module included in arotary wing vehicle showing that the servo module includes three servoactuators and three Z-links for varying the pitch of the upper and lowerrotors at different phase angles simultaneously;

FIG. 34 is an enlarged perspective view of two pitch controllerswashplates included in the servo module of FIG. 33 showing the pitchcontroller swashplates connected by a Z-link to actuate the swashplatesat different phase angles;

FIG. 35 is a plan view the swashplates and Z-link of FIG. 34 showing aswashplate phase angle of about 90 degrees;

FIG. 36 is an exploded assembly view of the Z-link pitch control linkageof FIGS. 33 and 34;

FIG. 37 is plan view of a rotary wing vehicle in accordance with thepresent disclosure showing an upper rotor phase angle (solid doublearrow) and a lower rotor phase angle (hollow double arrow) and aresulting total rotor system phase angle (combined solid and hollowdouble arrow);

FIG. 38 is a side elevation view of a rotorcraft power and controlsystem according to the current disclosure configured for an aircraftwith a single drive motor, two rotors and a pusher propeller;

FIG. 39 is an enlarged perspective view of the rotorcraft of FIG. 38showing details of the main shaft splitter and drive gears for thecounter-rotating rotors and the belt-drive system for the pusherpropeller;

FIG. 40 is a perspective end view of a main rotor mast configured withinternal passageways for a torque tube and electrical wiring orplumbing;

FIG. 41 is a perspective end view of a main rotor mast configured withinternal passageways for a torque tube and six mechanical sliderlinkages;

FIG. 42 is a perspective view of a main rotor mast assembly includingthe main rotor mast of FIG. 41 and six slider linkages engaging the sixinterior passageways and connected to upper and lower swashplates;

FIG. 43 is an enlarged perspective end view of the main rotor mastassembly of FIG. 42 showing six swashplate slider linkages engaging thesix interior mast passageways;

FIG. 44A is a perspective view of a slider linkage configured with adownward pointing follower link to control a lower swashplate;

FIG. 44B is an exploded perspective view of a slider linkage configuredwith an upward pointing follower link to control an upper swashplate;

FIG. 45 is a perspective side view of a helicopter with a non-rotatingmast and six rotary servo actuators coupled to the mast with upper andlower rotor hubs and rotor blades removed for clarity;

FIG. 46 is an enlarged perspective end view of the non-rotating mainrotor mast assembly of FIG. 45 showing the six rotary servo actuatorscoupled to the mast and connected to the upper and lower swashplateswith six individual linkages;

FIG. 47 is a perspective view of a high-speed helicopter in accordancewith the present disclosure showing that the high-speed helicopterincludes a non-rotating mast supporting an aerodynamic mask shroudbetween the upper and lower rotor blades to reduce drag;

FIG. 48 is an enlarged partial perspective side view of the helicopterof FIG. 47 with portions broken away to reveal the non-rotating mast,mast shroud, six linear servo actuators, and other control systemcomponents including electronics and antennae supported by the mastbetween the upper and lower rotor blades;

FIG. 49 is an enlarged partial perspective view of the non-rotating mastassembly of the helicopter FIG. 47 showing upper and lower rotor hubs,upper and lower rotor drive gears, and linear servo actuators;

FIG. 50 is a perspective view of the central non-rotating mast of themast assembly shown in FIG. 49 with the mast sleeve removed to showdetails of the electrical bus inlays;

FIG. 51 is a sectional view taken along line A-A of FIG. 50 showing atorque tube inside the mast and showing exterior channels for electricalbus inlays;

FIG. 52 is a perspective view of the electrical bus inlays of FIG. 51;

FIG. 53 is a perspective view of a mast sleeve with six interleavedlinear servo actuators and two swashplates configured to reduce a mastassembly frontal area; and

FIG. 54 is a plan view of the lower swashplate of FIG. 53 showing therelationship between the swashplate arms to reduce the frontal area ofthe mast assembly.

DETAILED DESCRIPTION

As suggested diagrammatically in FIG. 1, a rotary wing vehicle 1includes, in series, a first module 2, a first and a second rotor system3, 5, power modules 13 and 14, and a second module 15 coupled inspaced-apart relation to an airframe 40 extending along a common axis 7.Illustratively, airframe 40 is an elongated central backbone 40 and canbe arranged as a hollow core or having a cruciform cross-section. Inoperation, first rotor system 3, also called first rotor 3, and secondrotor system 5, also called second rotor 5, rotate in oppositedirections about common axis 7 to direct thrust in direction 24 andcreate lift in direction 24′ to cause controlled flight of rotary wingvehicle 1, as suggested in FIG. 2A. First module 2 is adapted to includea variety of guidance systems 50′, electronics 55, or payloads 15′.Second module 15 is adapted to include payload 15′, or in someembodiments, a variety of guidance systems 50′ and electronics systems55′. Payload 15′ may include, but is not limited to, munitions,radiation sensors, chemical detection sensors, biological agent sensors,active and passive listening devices, video sensors, supplemental powersources, or other mission-specific equipment. Rotary wing vehicle 1 thusprovides means for moving reconnaissance, observation, or surveymonitoring equipment to an area of interest to obtain informationtherefrom.

As suggested in FIGS. 1, 25A, and 25B, first rotor system 3 includes afirst motor 54, first rotor blades 20, and a first pitch controller 56.In illustrative embodiments, motor 54 is an electric motor as shown, forexample, in FIGS. 4-6, or other suitable means for providing power torotate rotor blades 20 about common axis 7. First rotor system 3 andsecond rotor system 5 are similar to one another in structure andfunction. Second rotor system 5 includes a second motor 61, second rotorblades 22, and a second pitch controller 57. In illustrativeembodiments, motor 61 is an electric motor as shown, for example, inFIGS. 4-6, or other suitable means for providing power to rotate rotorblades 22 about common axis 7. Illustratively, electrical and electroniccomponents are connected and communicate through electrical conduit 173and electronic conduit 174 which hold power and signal lines,respectively. Although rotary wing vehicle 1 is illustrated having tworotor systems, rotary wing vehicle 1 may have more than two rotorsystems as performance and mission demands dictate.

As shown in FIGS. 1 and 3, airframe 40 is non-rotating and forms acentral elongated hollow backbone to receive first module 2, first andsecond rotor systems 3, 5, power modules 13 and 14, and second module15. Illustratively, power modules 13 and 14 are positioned to lie inside-by-side relation to one another between second rotor system 5 andsecond module 15. Because airframe 40 is hollow power modules 13, 14 canbe connected electrically through the hollow backbone to motors 54 and61.

Illustratively, pitch controller 56 is a swashplate 56′ coupled to afore/aft servo 58 and a roll servo 59 to vary the cyclic pitch of rotorblades 20 in response to input from a controller 55. In someembodiments, swashplate 56′ is further coupled to a collective servo 98to collectively change the pitch of rotor blades 20. Likewise, pitchcontroller 57 is a swashplate 57′ coupled to a fore/aft servo 58 and aroll servo 59 to vary the cyclic pitch of rotor blades 20 in response toinput from a controller 55. In some embodiments, swashplate 57′ is alsocoupled to a collective servo 98 to collectively vary the pitch of rotorblades 20. In illustrative embodiments, controller 55 is a commandsignal controller as shown, for example, in FIG. 3, or other suitablemeans for providing a desired electrical or mechanical directionalsignal to servos 58, 59, or 98, and motors 54, 61.

Illustratively, rotary wing vehicle 1 has a fixed-pitch rotor systemhaving two servos 58, 59 for aircraft pitch (helicopter-style fore/aftcyclic input) or aircraft roll (helicopter-style right/left cyclicinput) control. Servo 98, shown in phantom in FIG. 1, can be mountedsimilarly to servos 58, 59 if collective pitch control is desired. Inembodiments having a fixed-pitch rotor system, rotor systems 3,5 areconnected to swashplates 56′, 57′ by pitch links 119. Servos 58, 59 areconnected to swashplates 56′, 57′ by pitch links 125, 126. A feature ofthe present disclosure is that rotary wing vehicle 1 can be flown withas few as one or two cyclic servo actuators (servo 58, 59). In a“one-servo” flight mode, differential torque of motors 54, 61 controlsyaw orientation, and servo 58 controls forward and backward flight. Withonly one cyclic servo, rotary wing vehicle 1, also called vehicle 1, canbe flown much like an airplane having only rudder and elevator control.In the illustrative “two-servo” flight mode, servos 58, 59 providefore/aft aircraft pitch and right/left aircraft roll control withdifferential torque of motors 54, 61 providing yaw control.

In operation, rotor hubs 101 rotate in opposite directions. Servos 58,59 are controlled by onboard flight control electronics to tiltsimultaneously swashplate 56′ and swashplate 57′ which then cyclicallyvary the blade pitch angle of rotating rotor blades 20 to tilt vehicle 1in one of aircraft pitch direction 170 and aircraft roll direction 171.In another embodiment having collective pitch (see FIG. 11B), collectiveservo 98 and a third pitch link (not shown) are provided to vary theaxial location of swashplates 56′, 57′ along common axis 7 and to varythe collective pitch of rotor blades 20, 22 using electronicCollective-Cyclic Pitch Mixing (CCPM). With collective-cyclic pitchmixing servos 58, 59, and 98 tilt swashplates 56′ and 57′ in unison tovary cyclic pitch and move swashplates 56′, 57′ axially in unison alongcommon axis 7 to vary collective pitch.

The illustrative embodiment employs differential motor speed for yaw(heading) control while in a vertical flight configuration. Normally,coaxial helicopters use variable blade pitch and differential bladeangle to control yaw motions in flight. In the present disclosure,differential torque generated by operating motors 54, 61 at differentspeeds relative to the fixed body of vehicle 1 generates yaw forces tostabilize and control yaw motion (i.e. rotation about common axis 7). Inthis method, the torque (and eventually the speed) of motor 54 isincreased or decreased in response to a yaw motion of rotary wingvehicle 1 about vertical common axis 7. The torque (speed) of secondmotor 61 is adjusted automatically by an onboard computer system,contained within controller 55, in opposition to the torque (speed) offirst motor 54 to maintain constant lift so that rotary wing vehicle 1neither gains nor loses altitude.

Rotor blades 20 and 22 are coupled to rotary wing vehicle 1, also calledrotary wing aircraft 1, and supported for rotation by rotor hubs 101.Rotor hubs 101 are further coupled for pivotable movement to an internalyolk 108, as shown best in FIG. 11A. Pivot axles 109 extend throughrotor hub 101 and are received by yolk 108. Yolk 108 is adapted tocouple a pair of rotor blades to rotor hub 101 for rotation about commonaxis 7. Yolk 108 is further coupled to a first end of a pair of pitchlinks 119. Each pitch link 119 is further coupled on a second end to aperimeter edge of swashplate 56′ or 57′. Thus, yolk 118 is pivoted byinput from swashplate 56′, 57′ in response to linear motion input fromservos 58, 59, or 98. This pivoting motion of yolk 118 in turn causeseach rotor blade 20, 22 to pivot in response, thus increasing ordecreasing the rotor blade pitch of rotor blades 20, 22.

As suggested in FIGS. 2A and 2B, a rotary wing vehicle 1 includes anupper section 2′, first and second rotors 3 and 5, a middle section 4, alower section 6, first and second power modules 13, 14, and a payload15′ arranged in spaced apart relation along common axis 7. Referring nowto FIGS. 2A-4, internal mechanical and electrical components withinupper section 2′ and middle section 4 of vehicle 1 are enclosed by athin-walled upper body shell 10 and a middle body shell 11,respectively. A lower body shell 12 covers a portion of lower section 6,but could be extended to cover all of lower section 6. A feature of thepresent disclosure is that body shells 10, 11 are blow-molded from aplastic material such as polycarbonate or ABS, and, in conjunction withbackbone 40, form a structure for rotary wing aircraft that has both acentral strength component and a thin exterior cover component thattogether are stiff, strong and easy to manufacture.

As shown in FIG. 3, a rotary wing aircraft 1 in accordance with thepresent disclosure has a rotor system comprising a motor 54 operablyconnected to rotor blades 20 by means of a drive train such as gears106, 107 (FIG. 11). A pitch control such as a swashplate 56′ (FIG. 10)is operably connected to rotor blades 20 to vary the cyclic and/orcollective pitch of rotor blades 20 in response to output from a servoactuator such as servos 58,59 (FIG. 3) through linkages such as pitchlinks 125, 126 (FIG. 10). Power such as electricity from batteries (notshown) or fuel from a storage tank (not shown) in a power module 13flows through a power conduit across rotor system and provides power tooperate controller 55, motor 54, and servos 58 and 59. Control signalsfrom controller 55 flow along a signal conduit and regulate the speed ofmotor 54 and the positioning output of servos 58 and 59. The powerconduit and signal conduit are conducted between an inflow side and anoutflow side of rotor blades 20 through channels 96, also calledinterior space 96, formed in the structural spine or backbone 40 (FIGS.7A, 7B, and 15) of vehicle 1.

In hovering flight, first rotor 3 and second rotor 5 rotate in oppositedirections about common axis 7 forcing air downward in direction 24 andlifting vehicle 1 in an upwardly direction, as suggested in FIG. 2A.First rotor 3 has rotor blades 20 configured to rotate in direction 21,and second rotor 5 has rotor blades 22 configured to rotate in direction23 about common axis 7. Because first rotor blades 20 and second rotorblades 22 are equipped with a cyclic pitch control, vehicle 1 isconfigured for directional flight in direction 25 wherein common axis 7is orientated substantially vertically.

Referring now to FIG. 2B, a second embodiment contemplated by thecurrent disclosure is depicted having a booster module 8 appended tolower section 6 at a booster interface 9. Booster module 8 contains, forexample, an auxiliary power source (not shown) to augment an internalpower source contained in power modules 13 and 14 carried in vehicle 1.Illustratively, the auxiliary power source (not shown) and power modules13 and 14 are electrical batteries 13 and 14. Booster module 8 includesleft and right wings 16, 17 to provide additional lift for vehicle 1 indirectional flight in direction 18 wherein common axis 7 is orientedsubstantially horizontally.

Airframe 40 forms a structural backbone of rotary wing vehicle 1 andgenerally runs vertically through the center of rotary wing vehicle 1from upper section 2′ to lower section 6, as shown best in FIG. 4.Illustratively, airframe 40 is a non-rotating core tube with a hollowinterior channel 96 (FIG. 7A) or a cruciform beam 97 with exteriorchannels (FIG. 7B). First and second rotor systems 3, 5, also calledfirst and second rotor modules 3, 5, all components within upper section2′, middle section 4, and lower section 6 are coupled to airframe 40.Referring now to FIG. 7A, elongated central backbone 40, also callednon-rotating hollow core tube 40, further acts as a conduit forelectrical wiring 45, plumbing (not shown), and mechanical linkages (notshown) passing between components in upper section 2′, middle section 4,and lower section 6 of rotary wing vehicle 1. Longitudinal slots 46 and47 are provided as entry and exits points for electrical wires 45,plumbing, and linkages. Since non-rotating hollow core tube 40 andcruciform beam are unitary and continuous between body sections 2, 4,and 6, the rigidity and light-weight structural properties of vehicle 1are increased. Illustratively, non-rotating hollow core tube 40 andcruciform beam 97 are preferably made of wound or pultruded carbongraphite fiber, fiberglass, or aluminum alloy number 7075 (or similar)with an outside diameter (core tube 40) or width dimension (cruciformbeam) of about 0.5 inches (13 mm) and a wall thickness of between about0.03 inches (0.76 mm) and about 0.05 inches (1.3 mm).

Rotary wing vehicle 1 is arranged having three body sections, as shownbest in FIG. 3. Upper section 2′ is arranged having a horizonsensor/stabilizer 50, an electronic gyro stabilizer 51, a gyro mountingtable 52 coupled to an upper end of core tube 40, a first motor speedcontroller 53, a first motor 54, a radio receiver, and controller 55.Middle section 4 includes a first swashplate 56′, a second swashplate57′, a fore-aft cyclic servo 58, and a roll cyclic servo 59. Lowersection 6 includes a second motor speed controller 60, a second motor61, a radio battery 62, first and second power modules 13 and 14, andpayload module 15.

In the illustrated embodiment, horizon sensor/stabilizer 50 is a model“FS8 Copilot” model by FMA company, electronic gyro stabilizer 51 is a“G500” model silicone ring gyro by JR company, motors 54, 61 are“B2041S” models by Hacker company, and motor speed controllers 53, 60are “Pegasus 35” models by Castle Creations company which arecomputer-based digital programmable speed controllers. Rotary wingvehicle 1 is also configured to receive a GPS receiver/controller andtelemetry system (not shown), arranged to be coupled to upper section2′.

Interior components of rotary wing vehicle 1 are coupled to core tube 40by ring mounts 70, as shown in FIG. 8. Ring mount 70 includes an annularinner portion 71 conforming to the annular exterior surface of core tube40. Ring mount 70 includes radially extending mounting arms 72, 73, 74having flanges 75, 76, 77 adapted to hold mechanical, electrical, andother interior components of rotary wing vehicle 1. Ring mount 70 isarranged to support motor 54 in flange 75, motor speed controller 53 onflange 76, and radio receiver 55″ on flange 77. Interior components ofvehicle 1 are coupled, for example, to mounting flanges using a varietyof fasteners (such a nylon ties through apertures 78) or adhesives.Annular portion 71 provides means for locking ring mount 70 tonon-rotating hollow core tube 40 to prevent ring mount 70 from rotatingor sliding axially along non-rotating hollow core tube 40. Means forlocking ring mount 70 to non-rotating hollow core tube 40 includesfasteners (not shown) received by set screw receiver 79 or a variety ofadhesives. A second ring mount 80, as shown in FIG. 9, includes anannular ring 63, arms 82 and 83, and axial posts 84, 85 for supportingbody standoffs 86, 87, 88, swashplate anti-rotation arms 90 and 91, andswashplate links 92 and 93.

Servo module 81 includes ring mount 80 supporting pitch servo 58, rollservo 59, and universal body standoffs 86, 87 (as described in U.S.Provisional Patent Application No. 60/525,585 to Arlton which is herebyincorporated by reference herein) which support middle body shell 11, asshown, for example, in FIG. 10. As suggested in FIGS. 3, 4, 5, 6, 9, 10and 15, body standoffs 86, 87, 88 are secured to ring mount 80.Through-holes 263 in body standoffs 86, 87, 88 are receptive to manytypes of commercial fasteners such as bolts and rods (not shown) forsecuring body standoffs 86, 87, 88 to ring mount 80 and middle bodyshell 11. Middle body shell 11 is generally secured to body standoffs86, 87, 88 to provide a cover and aerodynamic fairing for servos 58, 59and swashplates 56′, 57′. Ring mounts 70, 80 are arranged to incorporateand support many structural features of rotary wing vehicle 1. Ringmounts 70, 80 assist assembly of rotary wing vehicle 1 because ringmounts 70, 80 and associated interior components can be preassembled assubassemblies and then later assembled along with other modules tonon-rotating hollow core tube 40 in a final manufacturing step.

Referring now to FIGS. 11A, 12A, 12B, 13A, 13B and 14, rotor system 3,also called rotor module 3, includes a rotor mount 100, a rotor hub 101having an internal gear 107, first and second ball bearings 102 and 103,a shaft 101A extending between bearings 102 and 103, a ring clip 104,motor 54, a planetary gearbox 105, a pinion gear 106, a blade yolk 108,pivot axles 109, axle end caps 110, torsion springs 111, and rotorblades 20. A motor mount 122 is receptive to gearbox 105 to couple motor54 to rotor mount 100. When assembled, bearings 102, 103 are retained byring clip 104 engaging slot 99 on a boss 112 extending from rotor mount100. Rotor blade 20 is held in place by a pin 113 extending through cap110 and aperture 114 formed in axle 109. Axle 109 passes through abearing aperture 117 formed in rotor hub 101 and into an aperture 94 inyolk 108 when it is retained by another pin (not shown). Pitch links 119couple yolk 108 to swashplate 56′.

As shown in FIG. 11B, a rotor module adapted to support both cyclicallyand collectively pitchable rotor blades includes collective rotor hub201 that is similar to rotor hub 101 and receptive to a collective yolkframe 208 coupled to bosses 214 formed on an interior surface of hub 201by fasteners 212. Collective yolk frame 208 supports the radial flightloads produced by rotor blades 20 acting through thrust bearings 203.Pitch links 119 couple pitch arms 210 to swashplate 56′.

Illustratively, planetary gearbox 105 has a reducing speed ratio ofabout 4:1. Pinion gear on motor 54 has nine teeth and engages internalgear 107 on rotor hub 101 which has sixty teeth, so the total speedreduction ratio of rotor module 3 is about 26.7:1 (that is, the outputshaft of motor 54 turns 26.7 times for each turn of rotor hub 101). Thisreduction ratio encourages the use of high efficiency electric motorsrunning at high voltages and high speeds.

Illustratively, motor 54 is a brushless motor. In some applications,especially where flight times are short and economy is a factor (forexample, in a short-range disposable munition) several low-cost brushedmotors (i.e. motors having carbon brushes and rotating commutators) areused in place of one high-cost brushless motor 54 to turn rotor hub 101.In such cases, while rotor module 3 is shown having one motor 54 todrive rotor hub 101, it is within the scope of this disclosure toinclude several motors around the circumference of rotor mount 100 todrive rotor hub 101 instead of only one. It is also anticipated thatrotor hub 101 itself can be configured with wire coils and magnets toact as a motor so that no separate motors are required to drive rotorhub 101 about common axis 7.

Rotor blade 20 in the embodiment shown is injection molded ofpolycarbonate plastic material and is of the type described in U.S. Pat.No. 5,879,131 by Arlton, which patent is hereby incorporated byreference herein. Rotor blade 20 is free to flap upward and downwardabout 6 degrees about flapping axis 120 before tabs 121 on torsionsprings 111 contact pitch axle 109 and resist further flapping. Thismeans that rotor blades 20 can flap up and down freely in flight about+/−6 degrees and can fold upward 90 degrees and downward 90 degrees forstorage or during a crash landing.

In the embodiment shown in the drawings, rotor mount 100 is injectionmolded in one piece from a thermoplastic material such as polycarbonateor nylon. Rotor hub 101 is injection molded in one piece from athermoplastic material such as nylon or acetal. Rotor blades 20 aresupported in flight by rotor hub 101 (which forms part of the exteriorbody shell of vehicle 1 instead of by traditional coaxial shaftscoincident with common axis 7. This places rotor support bearings 102,103 very close to rotor blades 20 and frees space within the centralbody portion of rotary wing vehicle 1 for other mechanical or electricalcomponents. In a fixed-pitch rotor system (shown in the drawings) radialflight forces produced by rotating blades 20 are supported by internalyolk 108 which connects two rotor blades 20 and which includes aninternal aperture surrounding and bypassing core tube 40, thus nospecial thrust bearings are required.

Referring now to FIG. 15, a coaxial rotor system in accordance with thecurrent disclosure comprises core tube 40, two rotor systems 3, 5, twoswashplates 56′ and 57′, and one servo module 81 coupled to non-rotatinghollow core tube 40 in mirrored symmetry around servo module 81. While acoaxial rotor system with two rotors is disclosed, rotary wing vehicle 1could be equipped with additional rotor systems (not shown) spaced apartalong the length of non-rotating hollow core tube 40 for additionalthrust or operational capabilities.

In the illustrated embodiment, rotary wing vehicle 1 has a fixed-pitchrotor system which requires only two servos 58, 59 for aircraft pitch(fore-aft cyclic) and aircraft roll (right-left cyclic) control. A thirdcollective servo 98 can be mounted in a similar fashion in middlesection 4, for instance, if collective pitch control is desired.

Rotor systems 3,5 are connected to swashplates 56′, 57′ by pitch links119. Servos 58, 59 are connected to swashplates 56′, 57′ by pitch links125, 126. In operation, rotor hubs 101 rotate in opposite directions.Servos 58, 59 are controlled by onboard flight control electronics 55′to tilt simultaneously swashplate 56′ and swashplate 57′ which thencyclically vary the blade pitch angle of rotating rotor blades 20 totilt vehicle 1 in one of aircraft pitch direction and aircraft rolldirection. In another embodiment having collective pitch (see FIG. 11B), a third servo and third pitch link (not shown) are provided to varythe axial location of swashplates 56′, 57′ along common axis 7 and tovary the collective pitch of rotor blades 20, 22 using electronicCollective-Cyclic Pitch Mixing (CCPM). Using servos positioned to liebetween rotor systems 3, 5 and directly coupling control swashplates56′, 57′ with linkages to control a coaxial rotor system in this way isa feature of the embodiment.

An illustrative embodiment of the disclosure includes motors 54, 61positioned to lie above and below rotor blades 20, 22 (see FIG. 25A)with power transmission between the rotor systems 3, 5 accomplishedthrough electrical wiring 45 instead of mechanical shafting therebyreducing mechanical complexity and weight. In another embodiment (seeFIG. 26), motors 54, 61 are positioned to lie between the rotor blades20, 22, and servo actuators 58, 59 are positioned to lie in spaced-apartrelation to locate rotor blades 20, 22 therebetween (see FIG. 26).Because power and control of the rotor systems 3, 5 is entirelyelectrical in nature, the entire control system of rotary wing vehicle 1can be operated electrically by digital computers and solid-stateelectronics without mechanical linkages or hydraulic amplification.Locating the motors 54, 61, as shown in FIG. 25A, eliminates the needfor concentric rotating shafting between rotor blades 20, 22, andpositions servos 58, 59 to drive both swashplates 56′, (included infirst pitch controller 56) 57′ (included in second pitch controller 57)directly.

A feature of the present disclosure is that vehicle 1 can be flown withas few as one or two cyclic servo actuators (servo 58, 59). In aone-servo flight mode, differential torque of motors 54, 61 controls yaworientation, and servo 58 controls forward and backward flight. Withonly one cyclic servo, vehicle 1 can be flown much like an airplanehaving only rudder and elevator control. In a two-servo flight mode, asillustrated in the drawings, servos 58, 59 provide fore/aft aircraftpitch and right/left aircraft roll control with differential torque ofmotors 54, 61 providing yaw control.

In another embodiment of the current disclosure, power to drive motors54, 61 in flight is provided by high-capacity electric batteries 130such as lithium-polymer or lithium-ion batteries, or fuel cells.Referring now to FIGS. 16A and 16B, power module 13 has six rechargeablelithium ion batteries 130 arranged in a hexagonal pattern aroundnon-rotating hollow core tube 40 and wired in series to produce about21.6 volts of electrical potential. Battery ring mount 131 is formed toinclude center aperture (ring) 132 to accommodate non-rotating hollowcore tube 40 and flange 133 to hold batteries 130. Electrical wires 45from power module 13 enter non-rotating hollow core tube 40 at opening47 (see FIG. 7A), and are routed through non-rotating hollow core tube40 to motor speed controllers 53, 60.

As shown best in FIG. 25A multiple power modules 13, 14 are provided foradditional energy capacity during flight and are, illustratively, wiredin parallel to increase the electrical current available to motors 54,61. Flight times of rotary wing vehicle 1 can be adjusted by adjustingthe number of power modules 13, 14 carried in flight.

Extra locking rings (or ring mounts with no radial arms) 135 areprovided above and below power module 13, 14 to help couple powermodules 13, 14 to non-rotating hollow core tube 40, as shown, forexample, in FIG. 4. Since power modules 13, 14 are relatively heavycompared to other components of vehicle 1, locking rings 135 preventpower modules 13, 14 from sliding along non-rotating hollow core tube 40during a crash landing of rotary wing vehicle 1. A feature of thepresent disclosure is that rotary wing vehicle 1 is well-suited to bemanufactured and assembled in modules. Rotor, wing, control, power,booster, electronics, and payload modules are manufactured separatelyand slid onto core tube 40. Electrical connectors for connectionspassing through openings 46, 47 in core tube 40 are mounted flush withthe surface of core tube 40 to assist in assembly and disassembly ofvehicle 1 for maintenance and repairs.

Energy density and power density are considerations in UAV design andcan be applied to an aircraft as a whole. Aircraft with higher energydensities and power densities have better overall performance thanaircraft with lower densities. In general, energy density and powerdensity are defined as the amount of energy and power available per unitweight. For example, the energy density of a fuel or electric battery(also known as “specific energy”) corresponds to the amount of energycontained in a unit measure of fuel or battery (measured, for instance,in Nm/Kg or ft-lbs/slug).

Chemical (liquid) fuels tend to have higher energy densities thanelectric batteries. One additional characteristic of liquid fuel poweras compared to electric battery power is that the weight of a liquidfueled aircraft decreases over the course of a flight (as much as 60%)as it burns fuel. Consequently the energy density of a liquid fueledaircraft (i.e., the energy available per unit weight of the aircraft)decreases slowly and power density (power available per unit weight)increases as it flies. This means that the performance of liquid fueledaircraft actually improves near the end of a flight.

In contrast, the overall power density of an electric-powered aircraftis constant throughout the flight because the maximum output power ofthe batteries is almost constant and the batteries do not lose weight asthey discharge. Energy density also decreases quickly because the totalenergy available decreases. To improve energy and power density of thecurrent disclosure, an auxiliary electric booster or power module 8 isprovided that can be jettisoned in flight after its energy supply isdepleted. Thus, booster module 8 comprises additional battery modules(not shown) assembled around common axis 7 with a mechanism to retainbooster module 8 to rotary wing vehicle 1.

In another embodiment, booster module 8 includes an internal combustionengine (such as a diesel engine not shown) which drives an electricgenerator (not shown) to convert chemical energy contained in a chemicalfuel to electrical energy. In other embodiments contemplated by thisdisclosure, a turbo-electric generator system (not shown) may be used tocreate electrical energy. A consideration of a booster module 8containing such a gas-electric generator is that the entire weight ofthe module, fuel system, and engine, can be jettisoned at the end of afirst flight phase leaving the relatively low weight rotary wing vehicle1 to complete a second flight phase.

In the illustrative embodiment, booster module 8 includes foldable wings16, 17 to increase lift in a horizontal flight mode of rotary wingvehicle 1. As shown in FIG. 17, wing 17 is folded about folding axis 140for compact storage. Wings 16, 17 are attached at about their “quarterchord” location to pivot shafts (not shown). When deployed for flightwith pivot shafts held rigidly perpendicular to common axis 7 (see alsoFIG. 2), wing 16 is free to pivot about pitch axis 143 to find its ownbest angle of attack. Because wings 16, 17 are free to rotate abouttheir own pitch axes in flight, appendages such as wings 16, 17 aresometimes referred to as “free-wings.” It should be noted that wings 16,17, being free-wings, can operate efficiently over a wide speed rangebecause of their ability to change pitch automatically to meet theoncoming airflow. Application of such a free wing to a rotary wing UAVis a feature of the disclosure.

In high-speed horizontal flight, common axis 7 is orientatedsubstantially horizontally with rotor modules 3, 5 together acting likea single counter-rotating propeller to pull rotary wing vehicle 1 in ahorizontal direction 18. Wings 16, 17 help to lift lower section 6 andbooster module 8 so that rotor modules 3 and 5 can apply more power toforward propulsion and less to vertical lifting.

It should also be noted that the current disclosure does not requireaerodynamic control surfaces (such as on wings 16, 17) because cycliccontrol of rotor module 3, 5 provides control power for maneuvering inaircraft pitch (elevation) direction 144 and aircraft yaw (heading)direction 145 when common axis 7 is substantially horizontal.Airplane-style roll control (about common axis 7) during high-speedhorizontal flight is accomplished though differential torque/speed ofrotor modules 3, 5. This method of control for horizontal flight of arotary-wing UAV is a feature of the illustrative embodiment.

Referring now to FIGS. 18A and 18B, when the energy of booster module 8has been depleted, a command from on-board controller 55 of rotary wingvehicle 1 actuates a mechanism such as a latch (not shown) thatseparates booster module 8 from rotary wing vehicle 1 and booster module8 falls away in direction 19. Rotary wing vehicle 1 then, in one flightmode, assumes a more vertical orientation and flies like a helicopter.

In another embodiment, booster module 8 includes a mission-specificpayload 147 such as an explosive munition, dipping sonar, hydrophones,radio ID marker, or a sonobouy. As illustrated in FIG. 19, uponseparation from rotary wing vehicle 1, booster module 8 falls awayleaving a sonar or hydroponic system 147 or other sensor connected torotary wing vehicle 1 by wire or fiber optic cable 146 so that rotarywing vehicle 1 can move payload 147 from place to place, deliver payload147 accurately to a desired location, and act as a telemetry linkbetween payload 147 and a remote receiver (not shown). This can be aneffective method of, for example, monitoring a target or marking a shipat sea with a remote radio ID marker or other marking instrument.

FIG. 22 illustrates a method of delivering a marker comprising, forexample, a sensor, or a marking device, such as indelible paint or aradio transmitter, to a remote location, in this case a ship on an openocean 157. Vehicle 1 is shown approaching ship S (in frame), maneuveringto touch ship S and leaving the marker on ship S (in frame) and exitingthe area (in frame). This method of marking is a feature of the presentdisclosure that allows a point of interest to be monitored after vehicle1 has left the local area. Alternatively or in conjunction, vehicle 1can retain a sensor when it leaves the local area which may, forinstance, have taken a sample of the atmosphere near ship S, and returnthe sensor and sample to a remote processing point for further analysisby a mass spectrometer, biological or radiological measuring device orother such device (not shown). While the point of interest shown in thedrawings as a ship S, it will be understood that ship S could be anyother point of interest accessible to vehicle 1 such as a truck,aircraft, building, tower, power line, or open area of land.

Another embodiment of the current disclosure shown in FIGS. 20A, 20B,and 20C, has unequal length folding, coaxial rotor blades 148, 149 withupper blades 148 having a greater span than lower blades 149. This is afeature arranged so that during a crash landing of vehicle when upperblades 148 contact the ground 155 before lower, shorter blades 149 sothat upper blades 148 fold away from, or faster than, lower blades 149thereby reducing the possibility that upper blades 148 and lower blades149 will contact each other while still rotating at high speed. As shownin the drawings, lower blades 149 span about 20 to 22 inches (51 cm to56 cm).

The ability to fold for compact storage and for landing is anotherfeature of the current disclosure. As shown in FIGS. 21A and 21B, rotarywing vehicle 1 is compact enough to fit inside a standard A-sizesonobouy tube used by the United States Navy. The unique core-tubestructure of the current disclosure not only allows rotary wing vehicle1 to be miniaturized to fit within a sonobouy tube, it also absorbs theforces of launch with a Charge Actuated Device (CAD) from an aircraftsuch as the Navy's P-3 maritime surveillance aircraft.

In one embodiment suggested in FIG. 21A, disposable launch canister 150is provided to protect the aerodynamic surfaces of rotary wing vehicle 1as it is launched from an aircraft traveling 150-250 knots at analtitude of 10,000 to 20,000 feet. A parachute (not shown) attached tocanister 150 slows and stabilizes the descent of canister 150 whichseparates from rotary wing vehicle 1 at a lower altitude.Illustratively, rotary wing vehicle 1 is shown to scale and has a bodylength 30 of about 24 inches (51 cm), upper diameter 31 of about 2.25inches (5.7 cm), upper rotor diameter 32 of about 28 inches (71 cm) andlower rotor diameter 33 of about 24 inches (61 cm) or less. Boostermodule 8 has a length 34 of about 12 inches (30 cm). First rotor 3 andsecond rotor 5 rotate at about 1400 RPM in hovering flight and at aboutor above 2000 RPM during vertical ascent and high-speed maneuvers.

Another embodiment contemplated by this disclosure is adapted for usewith a munition for assessing target damage done by the munition. Asshown in FIG. 23, vehicle 1 is adapted for use with the munition,illustratively shown in the drawings as a gravity-delivered bomb 160.Bomb 160 is dropped from a launch platform such as an aircraft. Inoperation, gravity-delivered bomb 160 transports vehicle 1 to thevicinity of a target site whereupon vehicle 1 is released to fall awayfrom bomb 160, illustratively slowed by use of an auxiliary drag chute162, or ejected from bomb 160 by an explosive charge-actuated device,before bomb 160 reaches its target. Vehicle 1 then orbits or hovers inthe target area near the impact site to observe bomb damage andtransmits video and other information to a remote operator (not shown).This method of munition damage assessment is a feature of the disclosurewhich provides immediate battle damage assessments without requiring alaunch platform to remain in the strike zone and reduces the need forsubsequent strikes against the same target while minimizing risk tohuman crew members.

As shown in FIG. 26, motors 54, 61 are positioned to lie between rotorblades 20, 22. Servo actuators 58, 59 are positioned to lie inspaced-apart relation to locate rotor blades 20, 22 therebetween.

In another illustrative embodiment motors 54, 61 are located below rotorblades 22 and rotating torque tube 254 runs inside non-rotating masttube 253 for transmitting power to rotor 22 as shown, for example, inFIGS. 28-31. In another embodiment a gas engine (not shown) may beprovided to generate electric power from a heavy fuel such as dieselfuel or JP8 to operate motors 54, 61. In yet another embodiment, a gasengine (not shown) may be connected to torque tube 254 and rotor mount100 through a gearbox (not shown) to drive rotors blades 20, 22, alsocalled rotors 20, 22, about common axis 7, also called rotor axis 7.

Torque tube 254 may be connected directly to upper rotor hub 270 assuggested in FIGS. 28 and 29 or to a belt or gear powered transmissionand speed reduction system 271 provided at the upper end of mast tube253 as suggested in FIGS. 29 and 30. Speed reduction system 271, alsocalled transmission system 271, may be located at the upper end of masttube 253 so that torque tube 254 may be configured for high-speed,low-torque operation. As a result, torque tube 254 may be of lowerweight construction than a comparably sized main rotor shaft for ahelicopter that must support the full flight loads of rotor hub 270 andupper rotor blades 20.

Referring to FIGS. 27-31, rotary wing vehicles 250, 251 contemplated bythis disclosure include a streamlined body 260 and other featuressuitable for high-speed horizontal flight. Body 260 may be adapted insome embodiments to carry one or more human pilots or one or morepassenger. Rotary wing vehicles 250, 251 include counter-rotating rotorblades 20, 22 rotatable about common axis 7, landing gear 261,streamlined mast shroud 257, pusher propeller 258, and stabilizing tailfins 259. Mast shroud 257 is generally airfoiled in cross section whenviewed from above to reduce frontal drag. Mast shroud 257 is shownsecured to body shell 11 and hence by screws 277 to body shell standoffs86, 87, 88 which secure mast shroud 257 to mast tube 253 and preventmast shroud 257 from rotating about common axis 7.

As described in FIGS. 28 and 29, a rotor module 264 includes upper rotorblades 20, lower rotor blades 22, rotor control assembly 255, rotordrive assembly 262, and mast assembly 252. Rotor control assembly 255includes swashplates 56′, 57′, servos 58, 59, and pitch links 125, 126.Rotor drive assembly 262 includes motors 54, 61 with associated drivegears for driving rotors 20, 22 about rotor axis 7.

Mast assembly 252 includes torque tube 254 running inside mast tube 253and supported by upper mast bearing 273 and lower mast bearing 274 asshown in FIG. 32. Mast assembly 252 is secured to body 260 by mastbrackets 266, 267 and mast bolts 202.

Torque tube 254 is smaller in diameter than mast tube 253 leaving anannular space 275 running through the interior of mast tube 253 to actas a conduit for electrical wiring to servos 58, 59 and otherelectrical/electronic components. Wire slots 265, 269 are provided asentry and exits points for wiring, plumbing, and linkages (not shown).In one embodiment mast tube 253 is constructed of carbon fiber compositematerial and supports lateral flight loads produced by rotor blades 20,22 and damps in-flight vibration of torque tube 254 especially at uppermast bearing 273. Torque tube 254 may be constructed from carbon fiber,aluminum, or steel and may support vertical flight loads in addition totorsion. Mast bearing 273, 274 may be configured to support axial aswell as radial loads. Because mast tube 253 is generally rigid andnon-rotating, mast assembly 252 may be stronger and produce lessvibration than a rotor shaft on a conventional coaxial rotor helicopterwhich is generally unsupported by airframe structure above the lowerrotor.

Referring now to FIGS. 33-36, a rotor control assembly 282 in accordancewith one embodiment of the current disclosure includes upper swashplate279, lower swashplate 280, servo actuators 284, 285, 286, servo ringmounts 288, 289 and three blade pitch Z-links 291. While Z-link 291 maybe constructed as a single piece, it is shown in the drawings as anassembly of parts consisting of a generally rigid Z-link body 292 madeof glass-filled nylon and two wear-resistant universal ball links 293,294 made of a softer material such as unfilled nylon. Universal balllinks 293, 294 fit into link recesses 299, 300 in Z-link body 292 andare attached by screws 295.

Simultaneous, uniform, axial displacement of all three Z-links 291 inrotor control assembly 282, also called swashplate control assembly 282,parallel to common axis 7 causes swashplate 279 and swashplate 280 tomove axially along common axis 7 which displaces pitch links 119 therebychanging the collective pitch of rotor blades 20, 22 simultaneously.Non-uniform and independent axial displacement of Z-links 291 causesswashplates 279, 280 to tilt simultaneously inducing a cyclic pitchcontrol in rotor blades 20, 22. Z-links 291 are also constrained to moveparallel to common axis 7 by anti-rotation tabs 287 appended to ringmounts 288, 298 and act as swashplate anti-rotation links.

Z-link body 292 is configured to hold universal ball links 293, 294 at afixed differential phase angle 290 so that non-uniform axialdisplacement of Z-links 291 parallel to common axis 7 in direction 298causes swashplate 279 and swashplate 280 to tilt in different directionswhich affects the relative cyclic phase angle of rotor blades 20 and 22.Differential phase angle 290 is shown as 90 degrees but may lie betweenabout 60 to about 120 degrees depending on the characteristics of rotorblades 20, 22 and their speed of rotation. Differential phase angle 290may be changed by varying the length of universal ball links 293, 294.

Z-link 291 aligns the cyclic phase angles of upper rotor blades 20 andlower rotor blades 22. Rotor phase angle can be described as the anglemeasured between the cyclic pitch control input of a swashplate to arotor system of rotating rotor blades and the resulting flapping motionof the rotor blades and apparent tilt of the rotor disk. Normally thephase angle of a single rotor helicopter is close to 90 degrees.

Because of the aerodynamic interaction of the upper and lower blades ona coaxial rotor helicopter, however, the rotor phase response of eachrotor on a coaxial rotor helicopter is much different than 90 degrees.For instance as illustrated in FIG. 37, if upper swashplate 279 andlower swashplate 280 are tilted forward in direction 297, upper rotorblades 20 will appear to tilt in upper rotor phase direction 302 andlower rotor blades 22 will appear to tilt in lower rotor phase direction303 which means that the absolute upper and lower rotor phase angles areeach about 45 degrees. The phase angle difference 304 therefore is about90 degrees. When upper swashplate 279 and lower swashplate 280 are eachrotated 45 degrees about common axis 7 by the fixed differential phaseangle 290 of Z-links 291 before being tilted then upper rotor blades 20and lower rotor blades 22 will both appear to tilt in direction 297. Atthis point upper rotor blades 20 and lower rotor blades 22 are said tobe in phase with each other. Rotors that react in phase with each otherproduce powerful control forces.

As illustrated in FIGS. 38 and 39, a rotary wing vehicle according tothe current disclosure includes a streamlined fuselage or body 260, arotorcraft power and control system 306, a co-axial, counter-rotatingrotor system 307 capable of producing vertical lift and a rearwardfacing propeller 258 capable of producing horizontal thrust.

In operation, power from a motor or engine 309 turns first stage piniongear 311 which turns crown gear 312,313 in opposite directions asdescribed in FIGS. 38 and 39. Crown gear 312 is connected by a transfershaft to second stage pinion 314 which drives lower rotor main gear 316and lower rotors 22. Crown gear 313 is connected by a transfer shaft tosecond stage pinion 315 which drives upper rotor main gear 317, torquetube 254 inside mast 319 and upper rotors 20. A belt drive systemconsisting of pulleys 321,322 and V-belt 323 drive propeller shaft 324from the aft end of motor 309.

As illustrated in FIG. 40, a non-rotating structural mast 319 accordingto the current disclosure is configured with interior passageways orconduits 325 to accommodate both mechanical and electrical power andsignal transmission components. Mast 319 may include center column 326and outer sheath 327 which are generally circular in cross section andconnected by radially extending ribs 328 which function to both separateand stiffen center column 326 and outer sheath 327. In operation torquetube 254 runs between bearings 273, 274 (see FIG. 32) inside centercolumn 326 to transmit rotary motion from a power source located belowmast 319 to rotor blades 20 located near the upper end 318 of mast 319.Bearings 273, 274 act to align mast inside of center column 326 andprevent torque tube 254 from bending or touching the interior surface ofcenter column 326. Torque tube 254 is mechanically separated fromwiring, plumbing, hoses and linkages (not shown) which are locatedbetween center column 326 and outer sheath 327 in interior conduits 325.In essence, center column 326, outer sheath 327 and ribs 328 form aplurality of signal and power conduits which effectively separatemechanical, electrical and fluidic power and signal lines running insidemast 319.

Referring now to FIG. 41-43, a non-rotating structural mast 330according to the current disclosure is configured with six interiorpassageways 331 to accommodate swashplate linkages 332 that transfermechanical control signals from servos actuators (not shown) locatedbelow lower rotors 22 to swashplates 279, 280. Mast 330 may includecenter column 333 and outer sheath 334 which may be generally circularin cross section and connected by radially extending ribs 335 whichfunction to both separate and stiffen center column 333 and outer sheath334. In operation torque tube 254 runs inside center column 326 totransmit rotary motion from a power source located below rotor blades 22to rotor blades 20 located near the upper end 336 of mast 330.

Apertures or slots 342 may be provided in outer sheath 334 toaccommodate entry and exit of wiring, plumbing, hoses (not shown) andswashplate linkages 332. A feature of the current disclosure is thatribs 335 and center column 33 act to transmit structural loads aroundapertures 342 thereby improving the structural integrity of mast 330especially when many power and signal lines are routed through mast 330and much of outer sheath 334 is perforated by slots or holes. Anotherfeature is that apertures 342 may extend completely to an end 337 ofmast 330 to allow removal of mast 330 from an aircraft duringmaintenance operations. In one embodiment, power and signal linesrunning inside mast 330 may be removed and reinstalled without firstremoving plugs and connectors that may not easily fit through interiorpassageways 331 thereby reducing maintenance costs. Yet another featureof the current disclosure is that mast 330 may be economicallymanufactured, for instance, in an extrusion process from aluminum alloy7075 or in a pulltrusion process from epoxy impregnated carbon fibersfor low weight and high strength.

As shown in FIGS. 44A and 44B, each swashplate linkage 332 may beassembled from lower slider 338, upper slider 339, slider pushrod 340and pitch control link 341. Lower sliders 338 may be connected to aservo actuator (not shown) to move swashplate linkages 332 axiallyinside interior passageway 331 of mast 330. Upper sliders 339 arepivotably connected to pitch control links 341 which transmits axialmotion of swashplate linkages 332, also called swashplate sliders 332,to swasplates 279,280. Slider pushrod 340 is shown with threaded endsand rigidly connects upper slider 339 and lower slider 338 to move as aunit.

Three servo actuators (not shown) connected to lower sliders 338 maycooperate to move three swashplate linkages 332 to control upperswashplate 279 and the cyclic and collective pitch of rotor blades 20.Three additional servo actuators (not shown) connected to lower sliders338 may cooperate to move three swashplate linkages 332 to control lowerswashplate 280 and the cyclic and collective pitch of rotor blades 22.While shown in the drawings with pitch control link 341, swashplatelinkages 332 may also incorporate Z-link 291 in place of pitch controllink 341 in which case only three servos would be needed to control thecyclic and collective pitch of both rotor blades 20, 22.

As illustrated in FIGS. 45 and 46, a rotary wing vehicle 350 inaccordance with the present disclosure includes a streamlined fuselageor body 351, a co-axial, counter-rotating rotor system withcounter-rotating rotor blades (not shown) capable of producing verticallift and a rearward facing propeller 353 capable of producing horizontalthrust. A non-rotating backbone or mast 330 supports a plurality ofrotary output servo actuators 354 located behind mast 330 and aplurality of rotary output servo actuators 355 located in front of mast330. Servo actuators 354, 355 are configured to lie in close proximityto a longitudinally extending plane defined by common axis 7 andlongitudinal axis 356 to reduce the forward-facing surface area of theservo actuators 354, 355 in high-speed forward flight. This reduces thewidth of a shroud (not shown but similar to shroud 257 in FIG. 27 andshroud 368 shown in FIG. 48) needed to cover servo actuators 354, 355and minimize aerodynamic drag in high speed forward flight. Bolt holes357, as shown in FIG. 46, are provided to mount a streamlined mastshroud such as shroud 257. One feature of the current disclosure is thatcontrol system components such as servo actuators 354, 355 are locatedin front of and behind mast 330 to minimize the width of the mastassembly to reduce drag in forward flight.

Another embodiment of a rotary wing vehicle 360 is shown, for example inFIGS. 47-57. Rotary wing vehicle 360 includes a streamlined fuselage orbody 361, a co-axial, counter-rotating rotor system withcounter-rotating rotor blades 362, 375 capable of producing verticallift and a rearward facing propeller 353 capable of producing horizontalthrust. A non-rotating mast 364 supports mast sleeve 366 and a pluralityof linear (screwtype) servo actuators 365. In one example, the linear(screwtype) servo actuators 365 may be Moog model 880 Electric LinearServo Actuators that are mounted thereto by brackets or arms protrudingtherefrom. Servo actuators 365 are configured to lie in close proximityto a longitudinally extending plane defined by common axis 7 andlongitudinal axis 367 to reduce the width and aerodynamic drag of mastshroud 368 in high-speed forward flight. Engine 363, which may be a GET700 turboshaft engine for example, is provided to turn upper rotor 362about common axis 7 through gearbox 369, upper rotor drive gear 370 andupper rotor torque tube 379, and to turn lower rotor 375 through gearbox369 and lower rotor drive gear 371 attached to lower rotor shaft 380.

A feature of the current disclosure is that non-rotating mast 364 maysupport aircraft components inside of mast shroud 368 to take advantageof the air wake produced by mast shroud 368 in high-speed forwardflight. Electronic or hydraulic components 372, including, for example,hydraulic motors and hydraulic valves, and antennae 373 may be supportedby non-rotating bracket 374 in some embodiments. This reduces the needfor space inside the body 361, also called fuselage 361, of rotary wingvehicle 360 and places electronic or hydraulic components closer toservo actuators 365.

Non-rotating mast 364 may be fabricated from a metal or carbon fibercomposite material and include channels 376 extending axially along anexterior surface of mast 364 to accommodate electrical bus inlays 378 assuggested in FIGS. 50-52. Electrical bus inlays 378 extends from a point390 between upper and lower rotors 362, 375 to a point 391 below thelower rotor 375 and between upper rotor drive gear 370 and lower rotordrive gear 371 to facilitate transmission of electrical and/or hydraulicpower and signals from components located in fuselage 361 of rotary wingvehicle 360 to other components located between upper rotor 362 andlower rotor 375 or above the upper rotor 362. Electrical bus inlays 378may include a protective sheath made of a non-conducting material suchas silicone and contain a plurality of copper conductors or hoses 382.In one embodiment mast sleeve 366 slides over mast 364 to provide amounting structure for servo actuators 365 and bracket 374 and a smoothexterior running surface for swashplates 384,385. Apertures 387 may beprovided in mast sleeve 366 to provide access to copper conductors orhoses 382 for electrical or hydraulic connections (not shown) to othercomponents such as servo actuators 365 and flight control systemelectronics (not shown). In operation a plurality of electrical wiresand/or hydraulic hoses (not show for clarity) may connect to bus inlays378 at copper conductors or hoses 382 to transmit electrical orhydraulic power and signals to and from other control system componentssuch as a flight management system computer (not shown), servo drivers(not shown), hydraulic motor 372, hydraulic values (not shown), andgenerators (not shown). A sturdy truss structure 388 may be provided toconnect mast 364 to fuselage 361 of rotary wing vehicle 360.

An important feature of the current disclosure is the reduction ofaerodynamic drag in high-speed flight. To reduce the width andassociated drag of mast shroud 368, swashplates 384 and 385 areconfigured to locate all six servo actuators 365 in close proximity to alongitudinally extending plane defined by common axis 7 and longitudinalaxis 367 as illustrated in FIG. 53. Swashplate arms 392 and 393 arecloser to each other than arms 393 and 394. As shown in FIG. 54, angle395 is about 90 degrees or less. Swashplates 384 and 385 are alsorotated 180 degrees relative to each other about common axis 7 so thatservo actuators 365 may be interleaved around the circumference of mastsleeve 366 for a very compact installation.

One feature of the disclosure is the non-rotating hollow core tube 40,mast 330, 364 or cruciform beam structural backbone that can, in someembodiments, double as a conduit for wiring and plumbing. A method orsystem of assembling mechanical and electrical components to the core orbackbone is described to promote ease of assembly of a variety ofaircraft from a kit of basic modules.

Another feature is that each of the rotors 20, 22 of the coaxial systemof the current disclosure are driven by one or more separate electricmotors, and the motors are positioned to lie on opposites sides of therotors, with power transmission to and between the motors accomplishedthrough electrical wiring (passing through the hollow core) instead ofmechanical shafting, clutches, and gears. Compact rotor assembliessupport the rotors for rotation without the need for traditionalrotating coaxial shafting.

Still another feature is that a swashplate control system and one ormore electric motors may be provided for each rotor and may bepositioned to lie on opposite sides of each rotor thereby simplifyingthe mechanical and electrical connections needed to drive and controlthe rotors. Rotor modules are provided to quickly and easily assemblesystems of rotors to the hollow core. Multiple rotor modules andswashplates are controlled by a single group of servos housed in amodule.

Another feature of the disclosure is the provision of phase links toproduce differential phase control of the upper and lower rotorssimultaneously. In some embodiments, fixed-phase links can providecollective and cyclic control of both rotors with only three rotorcontrol servos instead of the four to six servos generally required forcoaxial rotor control.

Another feature is that full collective and cyclic control of the upperand lower rotor blades of a coaxial helicopter can be accomplished withservo actuators located below the lower rotor so that the axial distancebetween the upper and lower blades can be minimized.

Another feature is that a streamlined, non-rotating body shell may bemounted between the upper and lower rotor blades of a coaxial helicopterto reduce drag in high-speed forward flight.

Yet another feature of one embodiment is that power and control signalsmay be passed from a point located below the lower rotor blades to apoint located between the rotor blades to facilitate locating the rotorcontrol system, radio electronics, antennae, and other electrical andcontrol system components between the rotor blades to make productiveuse of the space between and the blades in high speed forward flight.

Yet another feature of one embodiment is that upper rotor blades 20 maybe driven by a torque tube 254 running inside the mast tube 253 andconnected to a motor 54 or engine located below rotor blades 22. Bothupper and lower rotors may be driven by a single gas-powered enginelocated below the rotors if desired.

An additional feature is that folding rotor blades 148, 149 are ofunequal length. On the current disclosure with counter-rotating rotors3, 5, folding blades 148, 149 of unequal length reduce the chance thatthe blades will contact one another as they fold at high speed during acrash-landing.

Another feature is that a mounting structure is provided betweencounter-rotating rotors 20, 22 to support a body shell 11 or other typeof aerodynamic fairing between rotor blades 20, 22. Body shell 11protects the control assembly 255 from weather and reduces the airresistance of exposed servos 58, 59, swashplates 56′, 57′, and pitchlinks 125, 126, also called pushrods 125, 126.

Another feature of the disclosure is a method of improving energy andpower density on UAV's which can include a booster module 8 which isseparable from the main vehicle in flight. A booster module 8 isprovided to operate the UAV during a first flight phase. At the end ofthe first flight phase, the booster module falls away thereby reducingthe weight of the UAV for continued operation in a second flight phase.On electric powered UAV's, the power module may comprise a pack ofbatteries with or without an auxiliary lifting surface which isjettisoned in flight after the battery power is depleted, or payloadsspecific to a particular mission.

1. A rotary wing aircraft comprising a first variable pitch rotor bladesupported for rotation about a common rotor axis of rotation in a firstrotor plane of rotation, a second variable pitch rotor blade supportedfor rotation about the common rotor axis of rotation in a second rotorplane of rotation, a first blade pitch controller located between thefirst rotor plane of rotation and the second rotor plane of rotation andconfigured to control a pitch of the first rotor blade, a second bladepitch controller located between the first rotor plane of rotation andthe second rotor plane of rotation and configured to control a pitch ofthe second rotor blades, and a first pitch control linkage having afirst end coupled to the first blade pitch controller and a second endcoupled to the second blade pitch controller so that displacement of thefirst pitch control linkage operates both the first blade pitchcontroller and second blade pitch controller at the same time.
 2. Therotary wing aircraft of claim 1, further comprising a first servoactuator located between the first rotor plane of rotation and thesecond rotor plane of rotation and coupled to the first pitch controllinkage to operate the first pitch control linkage.
 3. The rotary wingaircraft of claim 1, wherein displacement of the first pitch controllinkage causes the first blade pitch controller to vary a cyclic pitchof the first rotor blade and causes the second blade pitch controller tovary a cyclic pitch of the second rotor blade.
 4. The rotary wingaircraft of claim 1, wherein the first pitch control linkage isconfigured to tilt the first and second blade pitch controllers indifferent directions.
 5. The rotary wing aircraft of claim 4, whereinthe first pitch control linkage is a Z-link and the first end and secondend cooperate to hold the first blade pitch controller and second bladepitch controller at a fixed-phase angle relative to each other.
 6. Therotary wing aircraft of claim 4, wherein the fixed-phase angle between afirst blade pitch controller tilt axis and a second blade pitchcontroller tilt axis is between about 60 degrees and 120 degrees.
 7. Therotary wing aircraft of claim 1, further comprising a second pitchcontrol linkage interconnecting the first blade pitch controller and thesecond blade pitch controller to form a system of common pitch controllinkages and the system of common pitch control linkages cooperate tocontrol the cyclic pitch of the first and second rotor blades and acollective pitch of the first and second rotor blades.
 8. The rotarywing aircraft of claim 1, wherein displacement of the first pitchcontrol linkage causes the first blade pitch controller to vary thecyclic pitch of the first rotor blade and causes the second blade pitchcontroller to vary the cyclic pitch of the second rotor blades and acyclic phase angle of the first blade pitch controller is different froma cyclic phase angle of the second blade pitch controller.
 9. The rotarywing aircraft of claim 1, wherein the first pitch control linkage actsas an anti-rotation linkage and blocks the first and second blade pitchcontrollers from rotating about the common rotor axis of rotation.
 10. Arotary wing aircraft comprising a non-rotating structural backbone, afirst rotor system coupled to the non-rotating structural backbone andincluding a first variable pitch rotor blade supported for rotation in afirst rotor plane of rotation about an axis of rotation and a firstblade pitch controller, a second rotor system coupled to thenon-rotating structural backbone and including a second variable pitchrotor blade supported for rotation in a second rotor plane of rotationabout the axis of rotation and a second blade pitch controller, and afirst pitch linkage interconnecting the first blade pitch controller tothe second blade pitch controller, wherein the first blade pitchcontroller and the second blade pitch controller are positioned to liebetween the first rotor plane and second rotor plane in axiallyspaced-apart relation to one another and displacement of the first pitchlinkage operates the first blade pitch controller and second blade pitchcontroller simultaneously.
 11. The rotary wing aircraft of claim 10,further comprising a first servo actuator coupled to the non-rotatingstructural backbone and located between the first and second rotorplanes of rotation.
 12. The rotary wing aircraft of claim 10, furthercomprising an aerodynamic body shell supported by the non-rotatingstructural backbone and surrounding the first servo actuator and firstpitch control linkage to minimize air resistance of the rotary wingaircraft.
 13. A rotary wing aircraft comprising a first variable pitchrotor blade supported for rotation about a common rotor axis of rotationin a first rotor plane of rotation and the first variable pitch rotorblade includes cyclic pitch control, a second variable pitch rotor bladesupported for rotation about the common rotor axis of rotation in asecond rotor plane of rotation and the second variable pitch rotor bladeincludes cyclic pitch control, a first servo actuator located betweenthe first rotor plane of rotation and the second rotor plane of rotationand configured to vary a pitch of the first rotor blade, and anon-rotating structural backbone formed to include a generally hollowinterior extending along the common rotor axis of rotation through thesecond rotor plane of rotation and the first servo actuator is coupledto the non-rotating structural backbone.
 14. The rotary wing aircraft ofclaim 13, further comprising a power source located below the secondrotor plane of rotation and configured to provide power to drive thefirst variable pitch rotor blade and the power is transmitted throughthe non-rotating structural backbone.
 15. The rotary wing aircraft ofclaim 13, further comprising a first blade pitch controller locatedbetween the first rotor plane of rotation and the second rotor plane ofrotation for controlling a pitch of the first rotor blade and the firstservo actuator is coupled to the first blade pitch controller by a firstpitch linkage.
 16. The rotary wing aircraft of claim 15, furthercomprising a power source located below the second rotor plane ofrotation and configured to provide power to the first servo actuator andthe power is transmitted through the non-rotation structural backbone.17. The rotary wing aircraft of claim 15, further comprising a secondblade pitch controller configured to vary a pitch of the second variablepitch rotor blade and a second servo actuator supported by thenon-rotating structural backbone between the first rotor plane ofrotation and the second rotor plane of rotation and coupled to thesecond blade pitch controller by a second pitch linkage and wherein thesecond variable pitch rotor blade includes cyclic pitch control and thefirst servo actuator and the second servo actuator cooperate to vary thepitch of the first and second rotor blades.
 18. The rotary wing aircraftof claim 15, further comprising a second servo actuator coupled to thenon-rotating structural backbone between the first rotor plane ofrotation and the second rotor plane of rotation and a second pitchlinkage interconnecting the second servo actuator and the second bladepitch controller to vary a pitch of the second rotor blade.
 19. Therotary wing aircraft of claim 15, wherein the non-rotating structuralbackbone supports the first variable pitch rotor blade for rotationabout the common rotor axis of rotation.
 20. The rotary wing aircraft ofclaim 15, wherein power to drive the first variable pitch rotor bladeabout the common rotor axis of rotation is transmitted through a spaceformed in the non-rotating structural backbone.
 21. The rotary wingaircraft of claim 15, further comprising a second blade pitch controllerlocated between the first rotor plane of rotation and the second rotorplane of rotation and configured to control a pitch of the second rotorblade and the first pitch control linkage is coupled to the second bladepitch controller to cause the pitch of the first rotor blade and secondrotor blade to be varied simultaneously.
 22. The rotary wing aircraft ofclaim 15, further comprising an aerodynamic body shell coupled to thenon-rotating structural backbone and arranged to surround thenon-rotating structural backbone between the first rotor plane ofrotation and the second rotor plane of rotation.
 23. The rotary wingaircraft of claim 22, wherein the aerodynamic body shell is fixed to thenon-rotating structural backbone and is non-rotatable about the commonrotor axis of rotation.
 24. A rotary wing aircraft comprising anon-rotating structural backbone formed to include an interior spacetherein, a first variable pitch rotor blade supported for rotation abouta rotor axis of rotation in a first rotor plane of rotation and a firstswashplate connected to the non-rotating structural backbone forcontrolling a cyclic pitch of the first rotor blade, a second variablepitch rotor blade supported for rotation about the rotor axis ofrotation in a second rotor plane of rotation and a second swashplateconnected to the non-rotating structural backbone for controlling acyclic pitch of the second rotor blade, and a power module located belowthe second rotor plane of rotation and configured to provide power todrive the first variable pitch rotor blade about the rotor axis ofrotation, wherein the power is transmitted through the interior space.25. The rotary wing aircraft of claim 24, further comprising a firstmotor located above the second variable pitch rotor blade and coupled tothe first variable pitch rotor blade to drive the variable pitch rotorblade about the rotor axis of rotation and power is transmitted from thepower module to the first motor by wires running through the interiorspace of the non-rotating structural backbone.
 26. The rotary wingaircraft of claim 24, further comprising a first motor located below thesecond variable pitch rotor blade and coupled to the first variablepitch rotor blades by a rotating torque tube and the rotating torquetube is arranged to extend through the interior space of thenon-rotating structural backbone.
 27. The rotary wing aircraft of claim26, wherein the non-rotating structural backbone supports a majority offlight loads of the first variable pitch rotor blade and the rotatingtorque tube transmits primarily rotational motion to drive the firstvariable pitch rotor blade about the rotor axis of rotation.
 28. Therotary wing aircraft of claim 27, wherein the rotating torque tube issupported near the first variable pitch rotor blades by the non-rotatingstructural backbone.
 29. The rotary wing aircraft of claim 26, whereinthe non-rotating structural backbone supports horizontal flight loadsgenerated by the first variable pitch rotor blade and stabilizes therotation of the first variable pitch rotor blade about the rotor axis ofrotation.
 30. The rotary wing aircraft of claim 26, wherein thenon-rotating structural backbone supports vertical flight loadsgenerated by the first variable pitch rotor blade.
 31. The rotary wingaircraft of claim 26, wherein the rotating torque tube supports verticalflight loads generated by the first variable pitch rotor blade.
 32. Arotary wing aircraft comprising a non-rotating structural backboneformed to include a space therein, a first variable pitch rotor bladesupported for rotation about a rotor axis of rotation in a first rotorplane of rotation and a first swashplate connected to the non-rotatingstructural backbone for controlling a cyclic pitch of the first variablepitch rotor blades, a second variable pitch rotor blade supported forrotation about the rotor axis of rotation in a second rotor plane ofrotation and a second swashplate connected to the non-rotatingstructural backbone for controlling a cyclic pitch of the secondvariable pitch rotor blades, and a first motor located below the secondrotor plane of rotation and coupled to the first variable pitch rotorblade to drive the first variable pitch rotor blade about the rotor axisof rotation, wherein power to drive the first variable pitch rotor bladeabout the rotor axis of rotation is transmitted from the first motor tothe first variable pitch rotor blade by a rotating torque tube locatedin the space formed in the non-rotating structural backbone.
 33. Therotary wing aircraft of claim 32, further including a second motorlocated below the second rotor plane of rotation and coupled to thesecond variable pitch rotor blade to drive the second variable pitchrotor blade about the rotor axis of rotation.
 34. The rotary wingaircraft of claim 32, further comprising a speed reduction systemlocated near an upper end of the non-rotating structural backbone andarranged to interconnect the first variable pitch rotor blade and thetorque tube.
 35. The rotary wing aircraft of claim 32, furthercomprising an aircraft fuselage coupled to a lower end of thenon-rotating structural backbone.
 36. The rotary wing aircraft of claim32, wherein the non-rotating structural backbone supports flight loadsgenerated by the first variable pitch rotor blade and stabilizes therotation of the first variable pitch rotor blade about the rotor axis ofrotation.